GOE 758 AIRFOIL (goe758-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 758 AIRFOIL (goe758-il) Reynolds number: 1,000,000 Max Cl/Cd: 133.59 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe758-il-1000000.txt Download as CSV file: xf-goe758-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 758 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.1069 0.08443 0.08254 -0.0798 0.9014 0.0224
-10.750 -0.1025 0.08098 0.07899 -0.0812 0.8801 0.0233
-10.500 -0.1232 0.07143 0.06940 -0.0859 0.8703 0.0251
-10.250 -0.1329 0.06501 0.06292 -0.0883 0.8581 0.0251
-10.000 -0.1544 0.05735 0.05524 -0.0905 0.8467 0.0255
-9.750 -0.2874 0.03092 0.02856 -0.1106 0.8416 0.0256
-9.500 -0.2996 0.02904 0.02659 -0.1096 0.8269 0.0257
-9.250 -0.3128 0.02678 0.02419 -0.1078 0.8127 0.0258
-9.000 -0.3139 0.02458 0.02187 -0.1068 0.7985 0.0259
-8.750 -0.3048 0.02323 0.02043 -0.1059 0.7821 0.0261
-8.500 -0.3000 0.02127 0.01830 -0.1049 0.7631 0.0263
-8.250 -0.2900 0.01976 0.01664 -0.1040 0.7389 0.0267
-8.000 -0.2813 0.01791 0.01455 -0.1030 0.7117 0.0271
-7.750 -0.2699 0.01628 0.01267 -0.1019 0.6870 0.0279
-7.500 -0.2528 0.01658 0.01248 -0.1002 0.6690 0.0299
-7.250 -0.2536 0.01175 0.00711 -0.0988 0.6595 0.0305
-7.000 -0.2351 0.01013 0.00538 -0.0984 0.6503 0.0310
-6.750 -0.2130 0.00924 0.00444 -0.0981 0.6432 0.0314
-6.500 -0.1909 0.02020 0.01505 -0.1022 0.6448 0.0319
-6.250 -0.1721 0.01748 0.01187 -0.1005 0.6382 0.0281
-6.000 -0.1482 0.01662 0.01084 -0.0998 0.6323 0.0284
-5.750 -0.1237 0.01575 0.00983 -0.0993 0.6270 0.0285
-5.500 -0.0991 0.01493 0.00886 -0.0987 0.6218 0.0284
-5.250 -0.0738 0.01423 0.00805 -0.0982 0.6173 0.0284
-5.000 -0.0479 0.01357 0.00732 -0.0978 0.6132 0.0285
-4.750 -0.0221 0.01301 0.00668 -0.0974 0.6090 0.0285
-4.500 0.0037 0.01254 0.00613 -0.0969 0.6047 0.0287
-4.250 0.0298 0.01211 0.00565 -0.0965 0.6007 0.0288
-4.000 0.0566 0.01177 0.00528 -0.0962 0.5972 0.0293
-3.750 0.0835 0.01155 0.00501 -0.0960 0.5931 0.0297
-3.500 0.1098 0.01130 0.00470 -0.0956 0.5887 0.0299
-3.250 0.1361 0.01103 0.00439 -0.0952 0.5845 0.0301
-3.000 0.1615 0.01046 0.00382 -0.0947 0.5811 0.0304
-2.750 0.1874 0.01009 0.00343 -0.0943 0.5773 0.0307
-2.500 0.2134 0.00981 0.00312 -0.0940 0.5736 0.0311
-2.250 0.2396 0.00963 0.00289 -0.0936 0.5695 0.0316
-2.000 0.2668 0.00942 0.00269 -0.0934 0.5664 0.0321
-1.750 0.2942 0.00925 0.00251 -0.0932 0.5630 0.0326
-1.500 0.3214 0.00912 0.00236 -0.0930 0.5592 0.0335
-1.250 0.3484 0.00903 0.00223 -0.0928 0.5552 0.0345
-1.000 0.3757 0.00895 0.00213 -0.0926 0.5515 0.0352
-0.750 0.4035 0.00885 0.00204 -0.0925 0.5480 0.0358
-0.500 0.4309 0.00874 0.00191 -0.0923 0.5440 0.0375
-0.250 0.4581 0.00869 0.00183 -0.0921 0.5398 0.0396
0.000 0.4853 0.00865 0.00178 -0.0919 0.5357 0.0433
0.250 0.5113 0.00829 0.00172 -0.0916 0.5319 0.1416
0.500 0.5342 0.00764 0.00171 -0.0910 0.5276 0.3633
0.750 0.5581 0.00731 0.00176 -0.0904 0.5230 0.5070
1.000 0.5837 0.00716 0.00181 -0.0899 0.5186 0.5851
1.250 0.6097 0.00704 0.00185 -0.0895 0.5140 0.6423
1.500 0.6347 0.00694 0.00189 -0.0889 0.5091 0.6986
1.750 0.6576 0.00678 0.00195 -0.0878 0.5041 0.7767
2.000 0.7333 0.00648 0.00207 -0.0981 0.4968 0.9848
2.250 0.7779 0.00660 0.00213 -0.1019 0.4903 1.0000
2.500 0.8032 0.00666 0.00217 -0.1013 0.4849 1.0000
2.750 0.8280 0.00674 0.00221 -0.1007 0.4783 1.0000
3.000 0.8526 0.00684 0.00228 -0.1000 0.4719 1.0000
3.250 0.8777 0.00693 0.00234 -0.0994 0.4653 1.0000
3.500 0.9018 0.00706 0.00241 -0.0987 0.4584 1.0000
3.750 0.9271 0.00714 0.00249 -0.0982 0.4515 1.0000
4.000 0.9510 0.00729 0.00258 -0.0974 0.4430 1.0000
4.250 0.9758 0.00740 0.00267 -0.0968 0.4340 1.0000
4.500 0.9997 0.00756 0.00279 -0.0961 0.4248 1.0000
4.750 1.0233 0.00773 0.00290 -0.0953 0.4141 1.0000
5.000 1.0477 0.00787 0.00302 -0.0947 0.4057 1.0000
5.250 1.0710 0.00806 0.00317 -0.0938 0.3970 1.0000
5.500 1.0954 0.00820 0.00330 -0.0932 0.3892 1.0000
5.750 1.1186 0.00840 0.00346 -0.0924 0.3808 1.0000
6.000 1.1426 0.00856 0.00362 -0.0917 0.3727 1.0000
6.250 1.1655 0.00878 0.00379 -0.0909 0.3637 1.0000
6.500 1.1879 0.00901 0.00398 -0.0900 0.3532 1.0000
6.750 1.2113 0.00920 0.00416 -0.0892 0.3448 1.0000
7.000 1.2331 0.00946 0.00438 -0.0882 0.3351 1.0000
7.250 1.2556 0.00968 0.00458 -0.0874 0.3258 1.0000
7.500 1.2777 0.00992 0.00480 -0.0864 0.3180 1.0000
7.750 1.2988 0.01019 0.00504 -0.0853 0.3081 1.0000
8.000 1.3205 0.01043 0.00527 -0.0844 0.2990 1.0000
8.250 1.3405 0.01073 0.00554 -0.0831 0.2901 1.0000
8.500 1.3616 0.01097 0.00578 -0.0820 0.2812 1.0000
8.750 1.3808 0.01129 0.00607 -0.0806 0.2711 1.0000
9.000 1.3974 0.01166 0.00639 -0.0788 0.2578 1.0000
9.250 1.4110 0.01207 0.00674 -0.0764 0.2434 1.0000
9.500 1.4227 0.01254 0.00715 -0.0737 0.2269 1.0000
9.750 1.4286 0.01328 0.00774 -0.0701 0.2021 1.0000
10.000 1.4292 0.01428 0.00856 -0.0660 0.1696 1.0000
10.250 1.4179 0.01592 0.00991 -0.0605 0.1242 1.0000
10.500 1.4058 0.01783 0.01159 -0.0556 0.0864 1.0000
10.750 1.4086 0.01914 0.01285 -0.0529 0.0713 1.0000
11.000 1.4060 0.02091 0.01449 -0.0501 0.0486 1.0000
11.250 1.3973 0.02328 0.01673 -0.0471 0.0261 1.0000
11.500 1.4016 0.02491 0.01837 -0.0455 0.0214 1.0000
11.750 1.4076 0.02648 0.01997 -0.0442 0.0195 1.0000
12.000 1.4137 0.02812 0.02166 -0.0430 0.0181 1.0000
12.250 1.4215 0.02968 0.02328 -0.0421 0.0175 1.0000
12.500 1.4266 0.03154 0.02520 -0.0412 0.0166 1.0000
12.750 1.4298 0.03364 0.02736 -0.0404 0.0159 1.0000
13.000 1.4303 0.03610 0.02990 -0.0397 0.0152 1.0000
13.250 1.4332 0.03837 0.03225 -0.0391 0.0149 1.0000
13.500 1.4376 0.04054 0.03449 -0.0387 0.0146 1.0000
13.750 1.4399 0.04298 0.03700 -0.0383 0.0142 1.0000
14.000 1.4419 0.04546 0.03954 -0.0380 0.0136 1.0000
14.250 1.4422 0.04815 0.04230 -0.0377 0.0135 1.0000
14.500 1.4414 0.05099 0.04521 -0.0374 0.0130 1.0000
14.750 1.4378 0.05420 0.04849 -0.0372 0.0128 1.0000
15.000 1.4318 0.05771 0.05209 -0.0371 0.0125 1.0000
15.250 1.4174 0.06233 0.05681 -0.0371 0.0121 1.0000
15.500 1.4160 0.06547 0.06003 -0.0372 0.0120 1.0000
15.750 1.4148 0.06859 0.06323 -0.0373 0.0119 1.0000
16.000 1.4102 0.07220 0.06692 -0.0375 0.0118 1.0000
16.250 1.4085 0.07549 0.07029 -0.0378 0.0116 1.0000
16.500 1.4038 0.07924 0.07412 -0.0382 0.0114 1.0000
16.750 1.3996 0.08290 0.07786 -0.0386 0.0112 1.0000
17.000 1.3950 0.08669 0.08172 -0.0391 0.0110 1.0000
17.250 1.3907 0.09046 0.08557 -0.0397 0.0108 1.0000
17.500 1.3853 0.09447 0.08965 -0.0404 0.0106 1.0000
17.750 1.3794 0.09855 0.09381 -0.0412 0.0106 1.0000
18.000 1.3744 0.10254 0.09786 -0.0420 0.0104 1.0000
18.250 1.3697 0.10654 0.10192 -0.0430 0.0102 1.0000
18.500 1.3626 0.11097 0.10643 -0.0441 0.0101 1.0000
18.750 1.3552 0.11548 0.11099 -0.0454 0.0099 1.0000
19.000 1.3484 0.11987 0.11545 -0.0467 0.0098 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 758 AIRFOIL (goe758-il)